1. Field of the Invention
This invention relates generally to a deployable and tracked solar array mechanism for nano-satellites and, more particularly, the invention relates to a deployable and tracked solar array mechanism for nano-satellites for a wrap-folded solar wing formed with a frame that restrains the lengthwise edges of a center panel of the solar wing on each side prior to deployment.
2. Description of the Prior Art
CubeSats are small spacecraft often 10 cm×10 cm×30 cm. Existing CubeSat deployable solar array panels hinge from all four 30 cm bus faces in a maximum power configuration. While this yields the maximum area of cells, the structural architecture is not well suited to sun tracking and hence its Average Orbital Power (AOP) is severely limited due to non-optimal pointing during the majority of an orbit. The peak power of the existing State Of The Art (SOTA) CubeSat systems is 21 watts. When sun off pointing is factored in to obtain an Average Orbital Power (AOP), the power in a nadir pointing mission drops to 5 to 7 watts AOP.
In addition, restraining the solar array panels during liftoff, flight, and positioning of the spacecraft and deploying the solar array panels when ready presents numerous problems. It is of utmost importance to protect the solar array panels and to reliably release the solar array panels when desired. Due to the fact that the spacecraft is orbiting the Earth, it is not possible to “fix” any solar array panels or release mechanisms after the spacecraft is in orbit.